1. Field of the Invention
The present invention relates generally to materials and methods used to insulate structures from high temperatures and pressures generated during the combustion of fuels. The present invention is particularly suitable for insulating structures, including, but not limited to, dome structures, nozzle structures, and igniter structures of rocket motors, such as solid propellant rocket motors used in the aerospace industry.
2. State of the Art
Solid propellant rocket motors have a center bore and/or cavity in the aft end of the motor in which combustion products of the solid propellant flow and are directed through the throat of a nozzle. Combustion occurs on the surface of the propellant and the resulting combustion products, upon passing through the throat, expand and are expelled from the exit cone of the nozzle located at the aft-most end of the motor. Combustion products are accelerated from subsonic velocities at high pressure within the rocket motor to supersonic velocities at near ambient pressure as the combustion products pass through the exit cone of the nozzle. The very high velocities at which the combustion products have been accelerated and directed by way of the rocket motor provide the thrust needed to propel the craft, or spacecraft, to which the rocket motor or motors are mounted.
Open-ended solid propellant rocket motors typically have a much larger cavity in the aft end of the motor, referred to as the aft dome. The open-ended design is used to facilitate and ease the retraction of mold tooling used in forming the internal geometry of the propellant grain within the rocket motor. With open-ended rocket motors, combustion products can impinge directly on the aft dome at velocities exceeding 300 feet per second (91 m/s) before exiting the nozzle.
Because of the high temperatures and pressures at which propellant fuels burn, typically of the magnitude of 5000xc2x0 F. (2760xc2x0 C.) and 1500 PSI (10,341 KPa), it is necessary to provide the internal surface of such rocket motor domes, as well as other components and portions of the motor, with a thermally insulating material that can withstand the impingement of high-velocity gases and oxidized, or partially burned particles of fuel. The structure of the aft dome is typically made of aluminum, an alloy steel, or a fiber-resin composite and would quickly rupture if directly exposed to the high-velocity, high-temperature combustion gases and oxidized particulates. The insulating material also serves to contain and protect the immediately surrounding area of the motor from the large amount of heat generated by the rapid combustion of propellant fuels. Thus, the insulating material must not only be capable of withstanding the impact of high-velocity gases and particulates, which are very erosive to insulating materials, but must also be able to withstand being subjected to high temperatures and high pressures upon the firing of the rocket motor.
Rocket motors have a nozzle exit cone, which directs the burning gas out of the motor and away from the craft. Such exit cones can be a fixed-type cone, which is typically immovably mounted to the aft or rearward portion of the dome. Alternatively, and frequently, the exit cone can be a variable-angle or vectorable-type cone which is pivotably mounted to the aft portion of the dome so that the exit cone can be moved angularly within a selected range to vector, or steer the craft in which the motor is installed, thus providing more directional control of the vehicle. Typically, the exit cone of a vectorable-type nozzle can be vectored within a range of 0 degrees to 10 degrees. The exit cone, whether a fixed-type or a vectorable-type, is typically attached directly to the aft portion of the dome and is often canted at a preselected angle from the centerline, or longitudinal axis, of the motor. This is particularly true when the motor is configured as a strap-on booster rocket to provide increased launch capacity for a primary or core space vehicle. A cant angle of up to 10 degrees from the centerline, or longitudinal axis, of the motor is frequently used. However, other cant angles can be used as necessary. The cant is often necessary in crafts having multiple booster rockets, and is required to direct the exiting flame away from the centerline of the craft to prevent overheating or scorching of the craft itself or of adjacently mounted motors.
Thus, the dome of an open-ended rocket motor as well as the insulation contained within the interior of the dome must be configured so as to allow the exit cone of the motor to be canted at a preselected angle and/or vectorable within a preselected range of angles. The cant and/or sustained vector at a given angle gives rise to increased char in the aft dome as gases are turned to exit through the nozzle throat. The effect of the nozzle cant also results in higher manufacturing costs due to more complex machining, additional labor, and material scrap.
The art in the past utilized domes, typically made of a preselected metal alloy, in which two to three preformed rings of tape-wrapped carbon phenolic insulation were bonded into position in a consecutive fashion within the dome to form a thermally insulating barrier therein. Tape-wrapped carbon phenolic insulation has been used in the past to minimize inert weight due to increased thickness and because of its ability to withstand the mechanical and thermal erosion attributable to direct impingement of combustion products on the open-ended aft dome. Each of these rings were usually manufactured separately because of the complex geometry of sequentially increasing diameters in order to be fitted within the dome at a proper station. The hollow done like ewise increases in diameter as viewed from the aft position, or nozzle end, of the motor, and moving forward or away from the exit cone toward the dome where propellant fuel is undergoing combustion.
In order to better understand and appreciate the present invention, reference is made to exemplary prior art insulators installed in the domes of open-ended solid propellant rocket motors shown in FIGS. 1 and 2. FIG. 1 depicts an open-ended aft dome and nozzle of a motor having a fixed-type exit cone, whereas FIG. 2 depicts the open-ended dome and nozzle of a motor having a vectorable-type exit cone.
More particularly, motor 2 depicted in FIG. 1 is provided with a dome shell 10 which generally encases an open dome region 4 and a nozzle throat region 6, and an exit cone shell 22 which generally encases an exit cone region 8. Dome shell 10 is typically a shell made of a pre-selected metal alloy and includes a flange portion 26 for allowing the dome shell 10 to be sealed and secured to the motor chamber case (not shown). Exit cone shell 22 is also typically made of a metal alloy and exit cone insulative liner 24 is typically made of a tape-wrapped carbon fiber/phenolic composite material. The cant, or angle, at which the dome shell 10 must be configured in order to allow exit cone shell 22 to extend away from the horizontal longitudinal centerline of the motor is designated as angle xcex1. As mentioned previously, angle xcex1 can range from 0 to about 10 degrees, but xcex1 can be any suitable angle.
As can be seen in FIG. 1, nozzle throat region 6 is defined by an integral throat entry (ITE) 18 which is typically formed of a three-directional or four-directional tape-wrapped carbon-carbon composite material and is externally supported by a nozzle throat insulator 20 typically formed of a unidirectional tape-wrapped carbon fiber/phenolic material.
It can further be seen in FIG. 1 that open dome region 4 is defined by aft dome insulator 14 which lines the more aft portion of dome shell 10 back to the integral throat entry 18. Forward dome insulator 16 abuts with aft dome insulator 14 at joint interface 15 and insulatively lines dome shell 10 from joint interface 15 forward to flange portion 26. Located behind dome insulators 14 and 16 and thus between dome insulators 14 and 16 and the interior surface of dome shell 10 is a shear-ply layer 12. Shear-ply layer 12 is typically formed of an elastomeric material containing either silica powder or aramid fibers (e.g., fibers made of Kevlar(copyright) material) and a curable polymer such as ethylene propylene diene monomer (EPDM), which is commercially available from a number of sources. Shear-ply layer 12 provides a cushion between the somewhat rigid dome insulators 14 and 16 and the inner surface of dome shell 10 upon firing the rocket motor.
The construction of dome shell 10, shear-ply layer 12, aft dome insulator 14, and forward dome insulator 16 is generally as follows. Shear-ply layer 12 is typically hand laid into essentially the full inner surface of dome shell 10 by cutting and trimming calendered sheets into the proper size and configuration so as to conform tot he inner surface of dome shell 10. A bonding system such as Chemlok(copyright) 205 primer and Chemlok(copyright) 236 adhesive available from the Lord Corp. is used to ensure a proper bond between the shear-ply layer 12 and the inner surface of dome shell 10. Upon shear-ply layer 12 being properly laid and trimmed to fit the inner contour of dome shell 10, the shear-ply layer and dome shell are vacuum bagged and autoclaved to cure and bond shear-ply layer 12 onto the inner surface of dome shell 10. After being cured, the exposed surface of shear-ply layer 12 is machined to a final contour and surface finish suitable for accommodating the insulating material to be bonded thereto.
Dome insulators 14 and 16 are made in accordance with previously known materials and methods, and are first formed by laying down or wrapping a tape comprising carbonized rayon fibers that have been impregnated with, for example, a phenol formaldehyde resin about a work mandrel having a preselected contour and size which corresponds to the inner surface of the respective insulator. Such tape or tape wrap material is commonly referred to within the art as tape-wrapped carbon phenolic composite material. Such tape-wrapped carbon phenolic material is difficult to manufacture and is increasingly difficult to obtain commercially due to prior sources ceasing to manufacture the subcomponents of such material, or such sources no longer being in business. Such precursor tape was originally manufactured by North American Rayon Corp. (NARC), a fiber manufacturing subsidiary of North American Rockwell Corp. in which the subsidiary is no longer in business. The fiber tape upon having been carbonized and impregnated with phenolic resin (Fiberite product number MX 4926) sells for over US$100.00 per pound (US$220/Kg), but there is no U.S. supplier currently manufacturing such tape. A multiyear supply of the carbon fiber precursor was purchased and stockpiled by many companies when NARC announced their plans to cease production of the precursor fiber. Alternatively, rayon fiber is available from CYDSA Corporation, a Mexico-based supplier for which the mandatory qualifications and certifications have not yet been fully conducted in order to be approved as a certified vendor within the industry. However, once suitable and properly qualified rayon fiber has been obtained, the rayon fiber must first be woven into cloth or tape, and further must be carbonized by a skilled, and often difficult to locate, carbonizing facility prior to being impregnated with a phenol formaldehyde resin. Such a resin is available either from Borden Corp. as part no. SC1-008 or Ashland Chemical Corp. as part no. 91LD. One facility used in the past for impregnating the carbon fiber with phenolic resin has been Fiberite Corp of Winona, Minn. However, as a practical matter, it is very difficult to orchestrate and ensure that such carbon resin composite material is properly prepared to provide an end product of suitable quality.
Attempts at incorporating an alternate carbonized fiber precursor, polyacrylonitrile (PAN), have been unsuccessful, in spite of good erosion resistance and adequate thermal performance. The PAN fibers when woven into tape and impregnated with phenolic resin suffer from inferior mechanical properties (i.e., low inter-laminar cross-ply and shear strength).
After the carbon fiber has been carbonized and impregnated with a phenolic resin, it is initially wrapped about respectively sized mandrels to individually preform insulators 14 and 16. The mandrels carrying respective tape-wrapped insulator preforms are vacuum bagged and the insulator preforms are then autoclave cured. Upon being autoclave cured, or alternatively hydroclave cured, individually sized and configured insulators 14 and 16 are removed from their respective mandrels. Thereafter, the hollow, bowl-shaped insulators are machined in order to configure the back, or outer surface, of each insulator to have a contoured matching dome shell 10 with shear-ply layer 12 previously installed therein. Additionally, the joint interfaces, such as joint interfaces shown as 15 and 19 in FIG. 1, are machined so as to provide a proper surface for being cooperatively abutted against adjacent insulators within dome shell 10. This machining process is quite expensive in terms of machining time, associated skilled labor, and the amount of material to be removed and hence simply scrapped. This machining expense is especially amplified with respect to the construction of aft insulator 14 which must be extensively machined in order to provide the final contour needed to allow nozzle throat region 6 and exit cone region 8 to be canted at a selected angle xcex1. Upon examining the lower portion of open dome region 4 of FIG. 1, it can be seen that the larger diameter insulator 16 is generally symmetrical about the longitudinal centerline of the motor. However, smaller diameter insulator 14 is asymmetrical about the longitudinal centerline in order for the nozzle throat region 6 defined by integral throat entry 18 and exit cone region 8 defined by exit cone shell 22 and exit cone insulative liner 24 to extend downwardly to provide the required cant angle xcex1. This asymmetrical configuration of insulator 14 requires that an extensive amount of material be machined from the preform of insulator 14 because the preform may only be initially formed as a symmetrical workpiece in order for the resin-impregnated carbon fiber tape to be properly laid down on the mandrel. In other words, insulator 14 must first be made as a symmetrical, hollow bowl with the center missing, then a large portion of the backside of the preformed bowl must be removed (ranging upwards of 50% of the original material) by expensive and difficult multi-axis machining in order to provide the insulator having the necessary cant or angled configuration to match the interior of dome shell 10 and previously installed shear-ply layer 12.
Upon the respective joint interfaces and backsides of dome insulators 14 and 16 having been machined, insulator 14 is first installed and bonded into dome shell 10 against the inner facing surface of shear-ply layer 12 and longitudinally positioned against integral throat entry 18. A structural epoxy adhesive, such as EA-934NA or EA-9394 available from Hysol-Dexter, Pittsburgh, Calif., is typically used for such bonding of insulator 14 to shear-ply layer 12 and at the bonding interface between the aft edge of insulator 14 and forward edge of integral throat entry 18. Upon insulator 14 having been positioned into place within dome shell 10, larger diameter insulator 16 is then installed and bonded to the remaining exposed portion of shear-ply layer 12 and against the forward edge of aft insulator 14 to form a bonded joint interface 15, also referred to as a secondary bond line, between the two adjoining edges of insulators 14 and 16. Typically, the same epoxy adhesive, such as the previously mentioned EA-934NA or EA-9394, is also used as a bonding agent for the backside of bonding insulator 16 against the inwardly facing surface of shear-ply layer 12 and at the joint interface 15 between the aft edge of insulator 16 and the forward edge of insulator 14. Dome insulators 14 and 16 are fully cured and bonded to shear-ply layer 12, which was previously bonded into dome shell 10. Lastly, the inner surfaces of dome insulators 14 and 16 are machined to the final contour and surface finish which open dome region 4 is to have. Thereafter, the completed dome assembly is ready to be installed as a major subcomponent of a rocket motor to which other components can now be secured, including exit cone assembly 22/24.
Referring now to FIG. 2 of the drawings, illustrated is an open-ended rocket motor 32 having a vectorable-type, or movable, exit cone. Motor 32 has a dome region 34 defined by dome shell 40 having three dome insulators, referred to as aft insulator 44, middle insulator 46, and forward insulator 48. Motor 32 is further provided with an integral throat entry 50 and a throat support insulator 52. Located about the outer circumference of the throat region is a pivoting mechanism 58 which allows exit cone assembly 54/56 to be pivoted within a preselected range. As with motor 2, shown in FIG. 1, dome shell 40 having a flange portion 60 is typically formed of a metal alloy and is configured to allow nozzle throat region 36 and exit cone region 38 to be canted at a preselected angle xcex1 with respect to the longitudinal centerline of motor 32.
Other than there being three dome insulators 44, 46 and 48 positioned in an end-to-end consecutive manner, such as at joint interfaces 45 and 47, and the three dome insulators 44, 46 and 48 bonded against the inner surface of shear-ply layer 42 bonded within dome shell 40 of motor 32, as compared with only two dome insulators being bonded to a shear-ply layer 12 in the dome shell of motor 2, the previously discussed materials and procedures for constructing and installing shear-ply layer 42 and dome insulators 44, 46, and 48 within dome shell 40 are essentially the same as for shear-ply layer 12 and dome insulators 14 and 16. As with asymmetrical aft dome insulator 14 of the motor shown in FIG. 1, asymmetrical aft dome insulator 44 of the motor shown in FIG. 2 must also have extensive multi-axis machining performed thereon in order for insulator 44 to be properly configured to accommodate the cant angle of nozzle throat region 36 and exit cone region 38. However, because three dome insulators are preferred, if not required, for motors designed to have a canted and/or vectorable-type exit cone, the associated manufacturing, labor, and material scrap rate are thus increased proportionally.
In closed-ended rocket motors, the propellant fuel fills a majority of the aft dome cavity, thereby resulting in a xe2x80x9cclosed-endedxe2x80x9d motor geometry as compared with an xe2x80x9copen-endedxe2x80x9d motor geometry as previously discussed and shown in FIGS. 1 and 2. In closed-ended rocket motors, the propellant fuel is bonded directly to the insulation material and/or an elastomeric liner, or stress-relief flap, which is in turn bonded to the insulation material. A material used in the past for insulating chambers of rocket motors for forming internal insulation of structures incorporated in closed-ended rocket motors is set forth in Table 1-1 of an unclassified Department of the Navy data sheet entitled Molding Compound, Rubber, Butadiene Acrylonitrile, with Phenolic and Boric Acid, Compounding of, First Revision dated Oct. 13, 1966, and as changed on Sep. 21, 1967. For convenience, such table is set forth in Table 1:
Formulations for rubbers are typically called out in parts by weight (PBW). The vulcanizable rubber portion of the formulation (in this case the Hycar 1051) is arbitrarily given a PBW of 100 and all other ingredients are called out as PBW in a level relative to the 100 PBW of the vulcanizable rubber.
The above-listed ingredients needed to make the subject prior art molding compound are available commercially and the process of combining such ingredients is further set forth in the subject data sheet, incorporated herein by reference. Insulative material formed from the above molding compound was specifically designed, tested, and approved to be used as insulating material in closed-ended motors in which the subject insulating material needed particular characteristics compatible with the propellant fuel to be bonded thereto. Furthermore, it is known within the art that this molding compound has been used to form nozzle stationary shell insulators similar to insulator 182 in FIG. 6 and as indicated in Section II of an unclassified Department of the Navy document of the Poseidon C3 Propulsion Test Program (Document No. SH050-A2A01HTJ, Report No. 10 dated Jul. 1, 1971). The prior art material cited in Table 1 hereof was used as a stationary shell insulator for the Poseidon Second Stage Motor design.
Materials originally designed, tested, and approved for closed-ended motors, such as in Table 1 herein, have not generally been investigated for use in open-ended motors due to differing design constraints between closed-ended and open-ended motors. When incorporated in closed-ended rocket motor designs, the material in Table 1, herein, was subjected to tightly controlled pre-ignition environments and limited to motor ignition temperatures greater than 70xc2x0 F. (21xc2x0 C.) because of concerns with strain capability at low temperatures.
Open-ended motor firing environments for this invention range from 30xc2x0 F. to 100xc2x0 F. The severe thermo-mechanical erosive environments typical of the aft dome of an open-ended rocket motor result in increased insulation thickness and, hence, additional inert weight, in order to adequately protect the aft dome structure. For those elastomeric materials with superior erosion resistance, such as the material cited in Table 1 herein, problems with batch and product reproducibility and consistency, manufacturing defects (voids) and insulator cracking due to aging were deemed higher risk options and unacceptable for an open-ended rocket motor production program.
The following documents are exemplary of rocket motor insulators known within the art:
U.S. Pat. No. 4,492,779, issued to Junior et al. and entitled Aramid Polymer And Powder Filler Reinforced Elastomeric Composition For Use As A Rocket Motor Insulation, is directed to a process for insulating solid propellant rocket motors with a composition comprising aramid fibers, a powder filler, and vulcanizable elastomeric composition;
U.S. Pat. No. 4,600,732, issued to Junior et al. and entitled Polybenzimidazole Polymer And Powder Filler Reinforced Elastomeric Composition For Use As A Rocket Motor Insulation, is directed to an elastomeric composition comprising polybenzimidazole polymer fibers, a powder filler and a vulcanizable elastomeric composition;
U.S. Pat. No. 4,458,595, issued to Gerrish Jr. et al. and entitled Ablative Liner, is directed to an end-burning rocket motor having a first layer of silicone rubber and a second layer of an ablative lining placed between the rocket motor casing and the propellant grain; and
U.S. Pat. No. 4,956,397, issued to Rogowaski et al. entitled Insulating Liner For Solid Rocket Motor Containing Vulcanizable Elastomer And A Bond Promoter Which Is A novolac Epocy Or A Resole Treated Cellulose, is directed to an insulating liner for a solid rocket motor having a vulcanizable elastomeric composition, powder filler, and a cellulosic bond promotor.
Additionally, the inventors of the present invention are aware of the following U.S. patents:
U.S. Pat. No. 5,352,212, issued to Guillot and entitled Method of Insulating a Rocket Motor, is directed to compositions of insulations containing thermoplastic liquid crystal polymers, fibers, and particulate fillers;
U.S. Pat. No. 5,399,599, issued to Guillot and entitled Thermoplastic Elastomeric Internal Insulation for Rocket Motors for Low Temperature Applications, is directed to compositions of insulations containing thermoplastic elastomers, an inorganic phosphorus compound, a polyhydric alcohol, a silicone resin, and chopped fibers;
U.S. Pat. No. 5,498,649, issued to Guillot and entitled Low Density Thermoplastic Elastomeric Insulation for Rocket Motors, is directed to compositions of insulations containing thermoplastic elastomers, a maleic anhydride modified EPDM, and carefully selected fillers and chopped fibers; and
U.S. Pat. No. 5,762,746, issued to Hartwell, et al. and entitled Method of Internally Insulating a Propellant Combustion Chamber, is directed to compositions containing polyphosazene polymer and organic fiber filler.
Thus, it can be appreciated that there is a need within the art for an insulating material having ingredients that can be readily and economically obtained and that can be readily formed and bonded into selected structures of rocket motors more efficiently and more cost effectively, as compared to previously known materials and methods.
It can further be appreciated that there is a need within the art for an insulating material that can withstand the high temperatures, high pressures and high particulate velocities encountered in the burning of propellant fuels in rocket motors without a substantial burden in increased inert weight to protect dome structures and adjacent motor components and the craft.
Another need within the art is the ability to construct insulators which are to be positioned within selected structures and components of rocket motors, particularly those of asymmetric configuration, while minimizing the number and complexity of steps required to perform such insulators. Furthermore, there is a need to minimize the amount of scrap or wasted material in constructing such insulators as well as to minimize the amount of difficult and expensive multi-axis machining needed to construct such insulators.
There is yet a further need within the art for an insulating material having particular ablative qualities as well as having a coefficient of thermal expansion and strain modulus particularly suitable for thermally insulating certain components of open-ended rocket motors and which can be obtained at a significantly lower cost as compared to previously known insulative materials.
Another need within the art is for an insulating material that has favorable aging characteristics, i.e., once an insulator is constructed, it may be several years before the rocket motor in which the insulator is incorporated is actually fired. Thus the art would benefit from having insulating materials having improved aging characteristics.
Yet another need within the art is for an effective insulating material in which the compound forming such insulating material is easier to mix, has uniformly dispersed ingredients, and repeatedly provides insulative products of consistently high quality.
The present invention provides a method of insulating a structure of a rocket motor having at least one surface on which an ablative insulative material is to be disposed. The method includes disposing at least one layer of curable ablative insulative material on at least one surface of the selected structure. The insulative material is generally formed of a vulcanizable rubber, a flame retardant such as zinc borate, a phenolic resin and a cure system constituent, and may optionally have reinforcing fibers therein. The method preferably includes the insulative material being formed of a compound including, but not limited to, the following ingredients: acrylonitrile butadiene rubber, zinc borate, phenol formaldehyde resin, zinc oxide, tetramethyl thiuram disulfide, and stearic acid. The insulative material may optionally be provided with supportive or reinforcing fibers or fibrous elements such as aramid, cotton (cellulose), sisal, polybenzamidazole, mineral wool, nylon, polyester, or carbon fibers. The method also includes curing at least one layer of curable ablative insulative material that has been disposed on at least one surface.
The present invention additionally provides a composition for an ablative insulative elastomerized phenolic resin material generally comprising: vulcanizable rubber, a flame retardant such as zinc borate, a phenolic resin and a cure system constituent, and which may optionally have reinforcing fibers therein. Preferably, the composition comprises the following ingredients: acrylonitrile butadiene rubber; zinc borate; phenol formaldehyde resin; zinc oxide; tetramethyl thiuram disulfide; and stearic acid. Preferably, the individual ingredients have the following maximum parts by weight: acrylonitrile butadiene rubberxe2x80x94100; zinc boratexe2x80x9480; phenol formaldehyde resinxe2x80x94120; zinc oxidexe2x80x945; tetramethyl thiuram disulfidexe2x80x943; and stearic acidxe2x80x942. Preferably, a stoichiometric master batch is provided that consists of a mechanically ground and screened phenol formaldehyde resin and zinc borate wherein the zinc borate coats the ground resin and acts as a partitioning agent to prevent the ground resin from agglomerating into larger, nondispersable particles or clumps. The stoichiometric master batch offers enhanced repeatability and quality control in producing insulative materials from the disclosed composition by ensuring that the largest undispersed resin particle does not exceed 100 mesh in size.
The present invention further provides a method of constructing an internally insulated metal dome structure of an open-ended rocket motor. The method includes disposing a shear-ply layer formed of a preselected curable elastomeric material onto at least a portion of the inner surface of a dome structure and curing and bonding the shear-ply layer to a selected portion of the inner surface of the dome structure. The method further includes preforming at least one first dome insulator about a mandrel. This first dome insulator (or insulators) is preferably made of a carbon phenolic composite material comprising carbonized fibers impregnated with a preselected curable resin and generally has an outer surface and an inner surface. The first insulator is then precured. A carbon phenolic or other highly erosion-resistant and structurally stable materials is required adjacent to the integral throat entry (ITE) in order to ensure a smooth transition of combustion gases flowing into the nozzle throat region. The method further includes machining at least a portion of the first insulator to a final configuration and positioning and bonding the first insulator onto a selected portion of the inner surface of the shear-ply layer previously disposed and bonded onto the selected portion of the inner surface of the structure. The method yet further includes disposing at least one second dome insulator (or insulators) onto at least a portion of the inner surface of the dome structure longitudinally proximate to the first dome insulator, the second dome insulator being generally formed of a vulcanizable rubber, a flame retardant such as zinc borate, a phenolic resin and a cure system constituent, and which optionally may contain reinforcing fibers therein. Preferably, the second dome insulator is formed of a material formed of a composition comprising at least acrylonitrile butadiene rubber, zinc borate, a curable resin, tetramethyl thiuram disulfide, and stearic acid. Preferably, the composition is prepared by a mixing process incorporating a stoichiometric master batch of resin and zinc borate. The second dome insulator is then cured and bonded within the dome structure. Alternatively, the second dome insulator may be precured, machined to a final contour and bonded to the dome structure using an epoxy adhesive.
The present invention also provides a thermal barrier for thermally insulating a structure. The thermal barrier is positioned to insulate at least a predetermined portion of the structure and includes an ablative insulative material formed generally of a vulcanizable rubber, a flame retardant such as zinc borate, a phenolic resin and a cure system constituent, and which optionally may contain reinforcing fibers therein. Preferably, the ablative insulative material is formed of a curable compound including the following ingredients: acrylonitrile butadiene rubber, zinc borate, phenol formaldehyde resin, zinc oxide, tetramethyl thiuram disulfide, and stearic acid. The thermal barrier may also include a shear-ply layer formed of an elastomeric material positioned between the structure and a second insulative material formed of a fibrous material impregnated with a phenolic resin, wherein a portion of the ablative insulative material and a portion of the second insulative material abut against each other to form a secondary bond line therebetween.